Solid propellant rocket motor

ABSTRACT

A solid propellant rocket motor having a controlled rate of thrust buildup to a desired thrust level by the combined utilization of a regressive-burning controlled flow solid propellant igniter and a progressive-burning main solid propellant charge wherein the igniter is capable of operating in a vacuum and sustaining the burning of the propellant below its normal L* combustion limit until the burning propellant surface and motor chamber pressure has increased sufficiently to provide a stable motor chamber pressure.

ijite States Patent [1 1 Fletcher et al.

[ SOLID PROPELLANT ROCKET MOTOR [76] Inventors: James C. Fletcher,Administrator of the National Aeronautics and Space Administration withrespect to the invention of; Warren L. Dowler, Sierra Madre; John I.Shafer, Pasadena; John W. Behm, San Pedro; Leon D. Strand, Pasadena, allof Calif.

[22] Filed: May 28,1971 [21] Appl. No.: 147,996

[52] U.S. Cl. ..60/256, 60/254, 102/49.7, 1 102/49.8 [51 Int. Cl ..F02k9/04 [58] Field of Search ..60/256, 254, 39.82; 102/493, 49.7, 49.8, 103

[ 56] References Cited UNITED STATES PATENTS 3,065,597 11/1962 Adamsonetal ..60/254 451 May 1,1973

2,561,670 7/1951 Miller et al. ..60/256 3,250,829 5/ 1 966 Wall 102/1033,248,873 5/1966 Pase ..60/39.82 E

Primary Examiner-Carlton R. Croyle Assistant Examiner-Robert E. GarrettAttorney-Monte F. Mott, Wilfred Grifka and John R. Manning 57 ABSTRACT Asolid propellant rocket motor having a controlled rate of thrust buildupto a desired thrust level by the combined utilization of aregressive-burning controlled flow solid propellant igniter and aprogressiveburning main solid propellant charge wherein the igniter iscapable of operating in a vacuum and sustaining the burning of thepropellant below its normal L* combustion limit until the burningpropellant surface and motor chamber pressure has increased sufficientlyto provide a stable motor chamber pressure.

5 Claims, 5 Drawing Figures Patented May 1, 1973 3,729,935

2 Sheets-Sheet 2 FIG.5

300 IGNITER P Q a; x I l l l FIG.4

INVENTORS WARREN L. DOWLER JOHN I. SHAFER BY JOHN w. BEHM LEON o. STRA 0ATTORNEYS SOLID PROPELLANT ROCKET MOTOR ORIGIN OF THE INVENTION Theinvention described herein was made in the performance of work under aNASA contract and is subject to the provisions of Section 305 of theNational Aeronautics and Space Act of 1958, Public Law 83-568 (72 Stat.435; 42 USC 2457).

BACKGROUND OF THE INVENTION 1 Field of the Invention This invention isin the field of solid propellant rocket motors, more particularly theinvention is in the field of solid propellant motors having lowacceleration-rate ignition, useful where the vehicle cannot readilywithstand high acceleration transients.

2. Description of the Prior Art There is present interest directedtoward orbit insertion missions at Jupiter, Saturn and Mercury whichrequire low acceleration and low acceleration-rates. Thus only a lowthrust, long burning motor would be suitable for such a mission. Thecontemplated orbital spacecraft for such missions require long, highlyflexible appendages for the scientific instrumentation and for theradioisotope thermoelectric generator contemplated to be utilized. For avehicle sensitive to performance and weight, such appendages dictate amaximum spacecraft acceleration at about I g thus inherently requiringlow thrust and long burning times for outer planet orbit insertionmotors. Further, and of concern to the herein invention, the flexibleappendages necessitate low acceleration rates, referred to as g-Dots,during the starting and shut-down thrust transients in order to preventlimit cycling of the autopilot or damage to the equipment mounted on theappendages and the possibility of even snapping off or breaking theappendages. Although there have been recent missions which have had therequirement for the low acceleration rates, the requirements were metthrough the use of low thrust liquid propellant rockets. For the givenvehicle constraints, i.e., dimensional envelope and performance of thepropulsion system, solid propellant motors were unable to meet the lowacceleration and acceleration rates required without unacceptable weightpenalties. Typical solid propellant motors, when ignited, buildup thrustvery rapidly at the rate of 5-50g/second. The thrust buildup desired forthe presently discussed application, involves an ignition system that iscapable of building up thrust at a controlled rate of less than0.2g/second. In the past, there has been an attempt to control thrustbuildup, though not to the degree mentioned, nor for the same reason.

The past effort was basically designed to provide an overall control ofthrust during the operation of a solid propellant motor, not just in theignition phase of operation. The most common previous approach was toutilize a variable throat area nozzle which, in some instances, was madeof an ablative material or the use of a movable pintle nozzle where theeffect of throat area is controlled by the movement of the pintle. Theseapproaches are quite complex and costly. Further, the variable areaconcepts often pay a significant weight penalty due to the mechanicalstructural requirements to achieve the variation in the throat area. Thevariable throat area nozzle would be particularly utilized in the hereininvention together with an inhibited progressive burning surface of apropellant charge so as to maintain the motor chamber pressure above thelow-pressure combustion instability limit. Such a requirement has notpreviously existed and thus this technique had not been used, though theaforementioned problems of the variable area throat configurationsindicate that such would not be preferable for the herein invention.

Another means for achieving the herein results would involve mountingadditional small solid propellant motors onto the spacecraft to give aprecise time sequence of thrust. This is once again quite expensive anda weight and packaging penalty is encountered. Further, the ability toobtain the precisely timed sequence would be difficult, if notimpossible. As a result, this approach would require complex andexpensive means for obtaining the precisely timed sequence. Thus, it isdesirable to provide a simple, lightweight means for achieving acontrolled buildup of thrust in a solid propellant motor.

SUMMARY OF THE INVENTION Briefly, the herein invention comprises anignition system for a solid propellant rocket motor which is capable ofproviding thrust at a controlled rate even though operation of the motorbelow its normal L* extinction pressure is required. This is achieved byutilizing a progressive burning solid propellant charge within a motorcase. The charge or grain preferably is highly inhibited or restrictedsuch that the burning area is insufficient initially to provide a stablemotor pressure. Alternatively, a grain can be pre-designed to have aprogressive burn without being inhibited. Thus, the herein inventionprovides an igniter that will ignite under all back pressure conditionsand under low L* conditions. It ignites the nonrestricted portion of theinhibited propellant, for example, while providing the mass additionnecessary to sustain combustion until the burning area of the propellanthas increased sufficiently to provide a stable motor chamber pressure.The igniter utilized has a regressive long burning time grain design. Inother words, until the burning area of the main motor propellant grainis sufficient to produce a stable motor chamber pressure, thecontribution of the igniter gas to the main motor mass flow, issufficient to maintain a high enough overall motor pressure to sustainstable operation. By having an independent controlled rate of flow fromthe igniter, one can further control the rate at which the thrust levelis built up. In one embodiment of the invention, a plurality ofexternally mounted igniters can be utilized to achieve the purpose ofthis invention; while in another embodiment the igniter can be mountedwithin the chamber in the form of a torus surrounding the throat areainjecting a gas in a circular pattern onto the exposed inhibited grainsurface. The composition of the solid propellant grain and thecomposition of the solid igniter can be tailored or selected to achievethe desired results in accord with the methods set forth below and donot require new propellant formulations. The foregoing arrangement willprovide a controlled low rate of thrust buildup due to the effect of theigniter until chamber pressure sufficient to sustain burning of the mainmotor is reached, thus providing the desired end results sought by thisinvention. It is believed that the invention will be further understoodfrom the following detailed description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a typical cross sectionalplan view of the motor of this invention.

FIG. 2 is a fragmentary detail of the igniter of FIG. 1.

FIG. 3 is a typical cross sectional view of a second embodiment ofamotor in this invention.

FIG. 4 is a typical cross sectional view of the inhibited end of a mainsolid propellant charge of this invention.

FIG. 5 is a typical graph of the pressure versus time curves ofa motorof this invention.

DESCRIPTION OF THE PREFERRED EMBODIMENTS It is well known in solidpropellant technology that a solid propellant composition has an L*combustion limit. For a given propellant composition, experimentalstability data have been correlated with the motor L* and chamberpressure. L* is defined as the free volume of the combustion chamberdivided by the throat area of the nozzle. The motor L* limit is thepoint below which irregular combustion and extinguishment will occur fora given propellant. For a fixed throat area of the nozzle, the chamberpressure must exceed a certain limit (minimum) in order that stablecombustion will occur or, in other words, that the L* limit must besurpassed. In normal solid propellant motors, the L* limit is reachedalmost immediately upon ignition, that is, within miIli-seconds ofignition, such that immediately, stable combustion of the solidpropellant grain is achieved. This, however, entails an extremely rapidacceleration rate or buildup of thrust.

In practice, operation of a motor near its normal L* combustion limitproduces low frequency, low amplitude chamber pressure oscillation,characteristic of L* combustion instability near the extinction pressureof that motor. It is for this and other reasons that rocket motors havealways been quickly brought up to and maintained beyond the L* limit tooperate at low chamber pressure. If such oscillations can be toleratedby the vehicle, then it would be advantageous if one could ignite asolid propellant below the L* limit and maintain a main charge burning,while gradually bringing up the chamber pressure and thrust level toprevent the rapid acceleration. Concern about potentially unacceptableacceleration rates from oscillations produced in a rocket motoroperating below its L* limit, led to a series of computer runs utilizinga spacecraft model to determine the effect of the initial accelerationduring igniter initiation and acceleration rates during ignition on thespacecraft. It was found that a maximum initial acceleration of 0.4 gwas acceptable. An acceleration rate during igniter burning of 0.4g/second was also acceptable, whereas a maximum initial acceleration of0.6 g and an acceleration rate during igniter burning of 0.6 g/secondwas unacceptable. Of course, these figures will vary depending upongiven spacecraft mechanization and configuration. In view of the data,however, it was determined that a conservative value of 0.3 g maximuminitial acceleration and an acceleration rate during igniter burning of0.3 g/second could be adopted as design criteria for an ignition system.The maximum acceleration during normal burning of the motor was not toexceed 1.0 g.

Turning now to FIG. 1, there is seen a first embodiment of the hereininvention which comprises a motor chamber 11 having an insulation 13therein. An end burning main propellant charge 15 is disposed within thechamber 11 and bonded to the insulation 13 in a conventional manner bychemical reaction'between the insulator surface and the propellantduring propellant thermal curing. At an aft end 17 of the motor, thereis connected a fitting 19 which supports a motor nozzle 21. The nozzle21 can have a throat insert 23 of graphite or other suitable material.The construction of the motor casing and attachment of nozzle 21 theretois conventional in the art and does not form part of this invention.Surrounding the nozzle support element 19 within the chamber is a layerofinsulation 25 which can be of an ablative type material and serves toinsulate both the support 19 and portion of the nozzle from the heatgenerated within the chamber. Alternative to the restricted end burninggrain, a radial burning non-inhibited grain could be provided in orderto achieve the desired progressive burn.

In the embodiment of FIG. 1 an internal igniter 31 is utilized, mountedon the insulation layer 25 surrounding the nozzle support 19. Theinternal igniter 31 is thus in a form of a torus having an outer case 33of for example polycarbonate or other thermoplastic material having aplurality of nozzles 35 spaced about the case and firing radiallyoutwardly into the main chamber area 39. The nozzles 35 can be silicaphenolic or other similar material. A fragmentary detail of the igniteris seen in FIG. 2. A layer 37 of insulation coats the inside of theigniter shell 33 and bonds an igniter charge 40 thereto. The insulationcan be ofa conventional rubber composition utilized in igniter or solidpropellant technology.

FIG. 3 discloses an alternate embodiment of this invention having thesame overall construction as shown in FIG. 1 except that in place of thetorus igniter 31 mounted internally within the chamber, one or moreexternal igniters 41 can be utilized having an igniter charge 43 locatedtherein. The external igniter requires a thick layer of insulation 42 sothat it will not burn out prior to the main motor charge. The externaligniter 41 can be connected to the outer case 11 such that the gasestherefrom are directed through a nozzle 45 into the main chamber 39.

In both the embodiments the exposed end 47 of the main propellant charge15 is highly inhibited initially as seen in FIG. 4 wherein a pluralityof inhibitor strips 49 are radially disposed about the exposed face ofthe charge. The inhibiting strips 49 can, for example, be comprised ofablative elastomeric insulation material which is bonded to the face ofthe propellant in a normal manner utilized for inhibitors in solidpropellant technology.

Turning now to FIG. 5, there is seen a typical graph which explains theoperation of the motors of this invention. When the igniter is ignitedit acts as a small solid propellant motor that burns with sonic exhaustfor about 1% seconds of independent operation of, for example, its 2%second burning time, if such periods are selected. The combustion gasesin the torus version shown in FIG. 1 pass radially outward from thenozzles provided impinging on the dish-shaped propellant surface 47.Typical igniter pressure could decrease from 350 psi to about 180 psi.The decay is shown in FIG. 5. The main motor propellant surface which ishighly inhibited, to produce a highly progressive burning surface as thepropellant regresses under the inhibitor surface, would have a typicalburning curve as shown in the dotted line of FIG. 5, if it could burn ata very low pressure by itself where its initial pressure would be only 5to psia. In reality, because of the L* combustion limit of about 65 psiafor the system of this example, the motor would not burn by itself belowthat pressure. However, when the hot exhaust gases from the independentcontrolled flow igniter are injected into the main motor chamber 39, themass addition raises pressure to about 50 to 55 psi and the burning ofthe main charge is sustained below its L* limit by heat transfer andmass addition. The resultant low pressure and thrust level permits thespacecraft to meet its 0.3 g initial acceleration requirement. Thus, themotor will effectively be operating as indicated by the line markedmotor P which is the effective result of the contribution of the gasesfrom both the main motor and the igniter.

The main propellant burning surface, and consequently the chamberpressure, will increase with time in a controlled manner, until themotor is able to sustain combustion without mass addition from theigniter. The small thin inhibitor strips 49 are partially or completelyconsumed before being ejected out of the nozzle. The pressure-timerelationship shown in FIG. 5 represents nominal performance values. Inpractice, however, operation of the main motor below its normal L*combustion limit, has produced low frequency, low amplitude oscillationscharacteristic of L* combustion instability near the extinctionpressure.

The reason it is believed the herein inventive concept is successful isfirst related to the fact that the burning and stability of the mainmotor charge is directly affected by the pressure developed withinchamber 39 of the motor. In other words, the propellant combustion andthe gases generated by such burning from the main charge, is directlycoupled to the chamber 39 and there is thus the aforementionedinterdependence between the developed chamber pressure and stableburning of the main charge. On the other hand, the igniter is sodesigned that it will burn and continue burning independently of theback pressure developed in chamber 39 and is only dependent upon its ownchamber pressure within the igniter case itself. In other words, unlikethe main charge, there is no coupling effect due to the sonic flow inthe igniter nozzles, between the pressure in motor chamber 39 andburning performance of the igniter, so that the igniter can maintain gasand mass flow production regardless of what is occuring within thechamber 39. As shown in FIG. 5, the production of gas or mass flow fromthe main motor would not reach the L* limit until 1 second had elapsedin this embodiment. Yet through the contribution of the mass flow fromthe igniter, the motor has a P (chamber pressure) curve as shown suchthat it will operate at acceptably stable conditions even below the L*limit of the main motor charge for almost the first one-half second ofoperation. Since the major contribution to mass flow is derived from theigniter charge during initial acceleration over the first second ofoperation, during which period the main motor is not dependent upon itsown gas generating capacity to fully develop the desired main motorchamber pressure, one can thus control the rate of acceleration throughthe design of the igniter.

Experimental results have been used to determine the igniter mass flowrate/main motor mass flow rate ratio as a function of burning time toestablish minimum motor resultant pressure and thrust at the preferredmotor U value. The farther below the L* limit one wants to operate, thelarger the igniter contribution required. The amplitudes of the pressureoscillations that occur below the L" limit grow slowly with time,thereby increasing the oscillatory acceleration (g-dots). This wouldlimit how long one would want to burn below the L* limit.

Since it is contemplated that the herein motors will be utilized in thehard vacuum of outer space, ignition of the igniter is important. Theignition involves the addition of external heat to the igniterpropellant at a rate such that its own subsequent decomposition andcombustion produce sufficient heat to sustain further decomposition andburning when the external heat source is withdrawn. The process iscomplicated especially in a vacuum by the propellant low pressuredeflagration limit, that is the pressure below which the propellant willnot burn (exclusive ofigniter If limits), the "short residence times forigniter combustion gases and the rapid decrease in pressure that canquench the burning. As a result, a material such as Pyrofuze, made bySigmund-Cohn Co., which is a coaxial wire with an aluminum coresurrounded by palladium in intimate contact with the core, can be used.Pyrofuze has a property that upon being heated electrically orchemically to l,220F which is the melting point of aluminum, itinstantaneously generates a large exotherm under gas pressure or in avacuum clue to a resulting alloy. The temperature of the wire can reach5,000F, well above the 400 to 800F ignition temperature range of mostsolid propellants. The Pyrofuze wires can thus be imbedded into thepropellant charge 40 of the igniter composition within slots 51 providedtherein. Other conventional ignition materials can also be used in placeof the Pyrofuze.

When current is passed through the Pyrofuze, burning would be initiatedthroughout the charge circumference on the bottom of the slot so thathot combustion gases must pass over the walls of the slot to reach thenozzle, thus promoting flame spreading and tending to raise the pressurelocally near the line of ignition. Various other means can be devisedfor effectively igniting the igniter charge in a hard vacuum which donot form part of the herein invention.

EXAMPLE The test motor utilized to demonstrate the feasibility of thepresent invention had a 5 inch inside diameter and was 6 inches long.The propellant charge was cast and pressure cured in the motor chamber.The restrictor configuration utilized was cut from a Gen GARD V-52rubber made by General Tire and Rubber Company, that had been cured tothe desired thickness and bonded to the propellant surface with cement.The restrictor essentially covered all of the end of the grain leavingfour slots having a width of 0.14 inch, each slot passing through thecenter of the grain and extending within one-eighth inch of the outerdiameter. Thus, the

end surface of the propellant grain was severely inhibited. Thepropellant formulation was comprised of 64 weight percent ammoniumperchlorate, 16 weight percent aluminum, 16.3530 weight percent hydroxylterminated polypropylene oxide, 0.2946 weight percent Alrosperse, whichis a surface active agent made by Geigy Chemical Corporation, 0.231weight percent trimethylol propane, 0.4873 weight percent l-decanol,2.3463 weight percent 2-6 toluene diisocyanate, 0.0369 weight percentferric acetyl acetonate, and 0.2500 phenyl beta napthylamine. The lengthof the propellant grain was 1.5 inches, while the weight of thepropellant was 1.76 pounds. The restrictor had a thickness of 0.25 inch.Thus, the initial burning area of the propellant was 2.7 square incheswhich was equivalent to 14 percent of the total area of the end of thepropellant. The nozzle throat diameter of the motor was 0.432 inches.Thus, the initial ratio of propellant burning area to nozzle throat areawas 20. A single external igniter was utilized at a 3-inch insidediameter and was 4 inches long. The end-burning igniter charge was inturn ignited by a pyrotechnic paste initiator system. The formulation ofthe igniter charge was 78 weight percent ammonium perchlorate, 2 weightpercent aluminum, 17.3073 weight percent hydroxyl terminatedpolypropylene oxide, 0.4677 weight percent Alrosperse, 0.0554 weightpercent trimethylol propane, 1.8696 weight percent 2-6 toluenediisocyanate, 0.0500 weight percent ferric acetyl acetonate, 0.2500weight percent phenyl beta napthylamine.

The test was highly successful. Pressure traces of the igniter and mainmotor were obtained. Post-test inspection revealed restricter segmentswedged in the nozzle entrance, but no blocking of the nozzle. The mainmotor pressure exhibited low frequency oscillations, over a fairlydistinguishable interval of approximately 3.5 seconds. The oscillations,which appeared to begin at a chamber pressure of approximately 17 psia,grew in amplitude with increasing pressure, and reached a maximum of 19percent of the mean pressure at a pressure of approximately 28 psia;then rapidly dampened out and approached zero at a pressure ofapproximately 40 psia. The very low amplitude oscillations appeared tocontinue as the pressure increased, but were difficult to distinguishfrom the erratic pressure perturbations and instrumentation noise thatoccurred throughout the run. The frequency of the oscillations grew froman initial value of5 Hz (5 cycles per second) to a value of Hz at 30psia. At their maximum amplitude the pressure oscillations gave rates ofchange of pressure of 160-200 psi/s.

The fact that the main motor pressure began to oscillate independentlyin a nonacoustic fashion at a pressure of approximately 17 psiaindicates that ignition has occurred at or before this point. Theinitial U (free volume/nozzle throat area) of the main motor for thistest was approximately 350 inches. From the L*-motor stability data, thelow-pressure combustion limit for the propellant at an L* of 350 inchesis between 45 and 50 psia. It was concluded, therefore, that thefeasibility of the g-dot ignition concept had been demonstrated, in thata motor was ignited below its low-pressure L* combustion limit andsuccessfully made the transition to the stable operating region in acontrolled manner.

What is claimed is:

. A solid propellant motor capable of controlled accelerationcomprising:

a motor chamber, having a nozzle affixed thereto;

a main solid propellant grain within said chamber having a progressiveburning surface and an inhibited end burning design such that theburning area of the end of said grain during an initial ignitioninterval generates pressure below the extinction pressure of said motorchamber and would not initially sustain stable combustion when ignitedand said progressive burning surface when sufficiently ignited aftersaid interval provides a chamber pressure above the extinction pressurewhich is independently capable of sustaining stable combustion;

and at least one solid propellant igniter connected to said chamber,said igniter being capable of supplying sufficient mass flow to saidmotor chamber at a controlled rate for a period no longer than saidinterval to raise the chamber pressure above the extinction pressurewhereby said motor will maintain stable combustion conditions during andafter said interval.

2. The motor of claim 1 wherein:

said igniter is disposed externally of said motor chamber.

3. The motor of claim 1 wherein:

said igniter is disposed within said motor chamber.

4. The motor of claim 3 wherein:

said toroidal igniter has a plurality of nozzles therein for directingsaid gas to said grain.

5. A solid propellant motor capable of a controlled rate of accelerationcomprising:

a motor chamber, having a nozzle affixed thereto;

a main solid propellant grain within said chamber having a progressiveburning surface such that said grain would not initially sustain stablecombustion when ignited;

and at least one solid toroidal propellant igniter disposed in saidchamber for directing gas to said main propellant grain, said igniterbeing capable of supplying sufficient mass flow to said motor chamber ata controlled rate whereby said motor will maintain stable combustionconditions.

i i t I III

1. A solid propellant motor capable of controlled accelerationcomprising: a motor chamber, having a nozzle affixed thereto; a mainsolid propellant grain within said chamber having a progressive burningsurface and an inhibited end burning design such that the burning areaof the end of said grain during an initial ignition interval generatespressure below the extinction pressure of said motor chamber and wouldnot initially sustain stable combustion when ignited and saidprogressive burning surface when sufficiently ignited after saidinterval provides a chamber pressure above the extinction pressure whichis independently capable of sustaining stable combustion; and at leastone solid propellant igniter connected to said chamber, said igniterbeing capable of supplying sufficient mass flow to said motor chamber ata controlled rate for a period no longer than said interval to raise thechamber pressure above the extinction pressure whereby said motor willmaintain stable combustion conditions during and after said interval. 2.The motor of claim 1 wherein: said igniter is disposed externally ofsaid motor chamber.
 3. The motor of claim 1 wherein: said igniter isdisposed within said motor chamber.
 4. The motor of claim 3 wherein:said toroidal igniter has a plurality of nozzles therein for directingsaid gas to said grain.
 5. A solid propellant motor capable of acontrolled rate of acceleration comprising: a motor chamber, having anozzle affixed thereto; a main solid propellant grain within saidchamber having a progressive burning surface such that said grain wouldnot initially sustain stable combustion when ignited; and at least onesolid toroidal propellant igniter disposed in said chamber for directinggas to said main propellant grain, said igniter being capable ofsupplying sufficient mass flow to said motor chamber at a controlledrate whereby said motor will maintain stable combustion conditions.